1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to a turbine blade with an internal cooling air circuit.
2. Description of the Related Art including information disclosed under 37 CFR 1.97 and 1.98
A gas turbine engine includes a turbine section with a plurality of stages of turbine blades to extract mechanical energy from a hot gas flow produced within a combustor. The efficiency of the engine can be increased by passing a higher temperature flow into the turbine. Modern turbine airfoil materials allow for only a maximum temperature without damaging the airfoils. To increase the temperature of the flow, complex cooling circuitry have been proposed to provide for both internal convection cooling of the airfoils and film cooling to provide a layer of film cooling air over the external surface of the airfoil to add protection against the high temperature gas flow.
The efficiency of the engine can also be increased by minimizing the amount of cooling air used to cool these airfoils. Compressed air from the compressor of the engine is drawn off and passed through the turbine airfoils for use in cooling. Less less cooling air from the compressor increases the efficiency of the engine because more of the compressed air can be passed into the combustor to produce the hot gas flow. Thus, turbine airfoil designers attempt to minimize the amount of compressed cooling air used to cool the airfoils while providing the highest amount of cooling capability with the minimal amount of cooling air.
FIG. 1 shows the external pressure profile on a prior art turbine blade used in an aero engine. The highest external airfoil pressures are located on the pressure side of the blade as seen from the top line in FIG. 1. The suction side of the blade past the vertical dashed line in this figure shows a low pressure area.
FIG. 3 shows a prior art turbine blade having a 1+5+1 serpentine flow cooling circuit used to cool a first stage turbine blade. The cooling circuit includes a leading edge cooling supply channel 11 and a leading edge impingement channel 12 to provide cooling for the leading edge region of the blade. A showerhead arrangement of film cooling holes 13 connected to the impingement channel 12 provides film cooling for the leading edge surface of the blade.
The trailing edge region of the blade is cooled by a circuit that includes a trailing edge cooling supply channel 15 with double impingement cooling channels 16 and 17 located downstream from the supply channel 15. Film cooling holes 18 connected to the supply channel 15 and trailing edge exit holes or slots 19 discharge cooling air from the impingement channel 17 out through the trailing edge of the blade.
The area of the blade between the leading edge and the trailing edge regions is cooled by a 5-pass serpentine flow cooling circuit that flows in the forward direction. This 5-pass serpentine flow circuit includes a first leg 21 that is the cooling supply channel for the 5-pass circuit, a second leg 22, a third leg 23, a fourth leg 24 and a fifth leg 25 that flows in series along the serpentine flow path from the leading edge end to the trailing edge end. The first leg 21 is an up-pass channel and includes a row of film cooling holes discharging on the pressure side of the blade. The second leg 22 is a down pass channel and includes a row of film cooling holes discharging on the pressure side of the blade. The third leg 23 is an up-pass channel, and includes two rows of film cooling holes to discharge film cooling air onto the pressure side and the suction side of the blade. The fourth leg is a down-pass leg, and includes a row of film cooling holes to discharge cooling air onto the pressure side of the blade. The fifth leg is an up-pass channel, and includes two rows of film cooling holes to discharge film cooling air onto the pressure side and the suction side of the blade.
In the prior art turbine blade cooling circuit of FIG. 3, cooling air is supplied to three separate channels of the blade as seen by the diagram of FIG. 4 representing the cooling flow paths. This forward flowing 5-pass serpentine circuit is used in the airfoil mid-chord region. The cooling air flows in the forward direction (from leading edge to trailing edge) and discharges into the high hot gas side pressure section of the pressure side. In order to satisfy the back flow margin criteria, a high cooling supply pressure is needed for this particular design, and thus inducing a high leakage flow. If a single channel includes film cooling holes on both sides of the airfoil, such as the third and fifth leg channels 23 and 25, and because the external pressure profile on the suction side has a lower pressure than on the pressure side, an excess pressure ratio across the suction side row of film cooling holes is developed.
Because the second and third up-pass channels (channels 23 and 25) of the 5-pass serpentine flow circuit provides film cooling air for both sides of the blade, in order to satisfy the back flow margin criteria for the pressure side film row, the internal cavity or channel pressure has to be approximately 10% higher than the pressure side hot gas side pressure which will result in over-pressuring the airfoil suction side film holes.
U.S. Pat. No. 5,813,835 issued to Corsmeier et al Sep. 29, 1998 and entitled AIR-COOLED TURBINE BLADE discloses an air cooled gas turbine blade with one serpentine cooling passage on a pressure side of the airfoil, a second serpentine passage on the suction side, and a third serpentine passage disposed in the middle of the airfoil. Film cooling holes on the last leg of the passages discharge cooling air from the blade. The present invention differs from the Corsmeier patent in that the three serpentine passages in the present invention all have a first leg on the pressure side of the blade, and also the second legs of the passages move to the other side of the blade on the suction side.
The U.S. Pat. No. 5,538,394 issued to Inomata et al on Jul. 23, 1996 and entitled COOLED TURBINE BLADE FOR A GAS TURBINE discloses (in FIG. 1 of Inomata) a turbine blade with three separate serpentine flow passages or circuits within the blade. A first 3-pass serpentine circuit includes a first leg on the pressure side, and a second and third leg on the suction side and adjacent to each other. Film cooling holes connect the third leg to the suction side surface of the blade. The second serpentine circuit is a 5-pass serpentine circuit having a first leg on the pressure side, a second leg on the suction side, a third leg on the pressure side, a fourth leg on the suction side, and a fifth leg on the suction side and adjacent to the fourth leg. Film cooling holes are connected to the fifth leg to discharge the cooling air. The third serpentine circuit is in the trailing edge region and includes a 3-pass serpentine circuit with each channel or leg extending between both the pressure side and suction side walls such that the legs do not alternate between sides. A separate leading edge cooling supply channel supplies cooling air to the leading edge cooling cavity and showerhead holes.
The Inomata cooling circuit requires more flow than the present invention because separate supply channels are required for the leading edge cooling circuit and the first serpentine circuit in the mid-chord region. In the cooling circuit of the present invention, the first leg of the serpentine circuit also supplies cooling air to the leading edge cooling cavity and showerhead holes. In the Inomata cooling circuit, the cooling supply channel for the leading edge region would produce a low mach number in the flow because of the narrowing channel toward the blade tip and the loss of flow as cooling air is metered off through the metering holes and into the leading edge cavity and showerhead holes. Also, the leading edge supply channel extends between both the pressure side and the suction side walls, and film cooling holes are located on both sides of the channel. The same pressure exists within the channel to discharge cooling air through the suction side film cooling holes as does the pressure side film cooling holes. In the cooling circuit of the present invention, the pressure side film cooling holes are connected to a different channel than are the suction side film cooling holes.
It is an object of the present invention to provide for a turbine blade with a serpentine flow cooling circuit that will optimize the use of the main stream pressure gradient.
It is another object of the present invention to provide a turbine blade with three separate 3-pass serpentine aft flowing cooling circuits.